CMC blade with integral 3D woven platform

ABSTRACT

A method of forming a component for use in a gas turbine engine includes the steps of forming an airfoil/root assembly; creating a platform assembly structure having an opening; inserting the airfoil/root assembly into the opening; and bonding the platform assembly structure to the airfoil/root assembly to form the component.

BACKGROUND

The present invention relates to a component, such as a blade, having anintegral platform for use in a gas turbine engine and a method forforming same.

Turbine blades typically require integral platforms, attached to theroot region, to form the inner flowpath, and to protect the rim of thedisk from hot gas ingestion. The typical cantilevered platform often hasto support Blade-to-blade dampers. The combined load of the overhungplatform and damper results in large bending stresses at the point wherethe platform meets the blade root region.

Ceramic Matrix Composites (CMC) are desired for turbine blades due totheir high temperature capability. CMC blades made from layers of clothor unidirectional tape offer good strength in the primary radial loadpath. Attaching a platform onto a laminated blade is difficult, and madeespecially challenging due to the low interlaminar strengths of CMC's.The bending of the platform causes high internal interlaminar stresses,and is a limiter on how long the platforms can function.

SUMMARY

In accordance with the instant disclosure, there is provided a method offorming a component for use in a gas turbine engine broadly comprisingthe steps of: forming an airfoil/root assembly; creating a platformassembly structure having an opening; inserting said airfoil/rootassembly into said opening; and bonding said platform assembly structureto said airfoil/root assembly to form said component.

Further, in accordance with the instant disclosure, there is provided acomponent for use in a gas turbine engine which broadly comprises anairfoil and a root portion formed from a ceramic matrix composite and athree dimensional platform assembly bonded to said airfoil and rootportion. The purpose is to create a platform region, with the inherentcomplexities, which is independent of the airfoil root structure, thusisolating the large radial pull of the airfoil from the platformassembly. Additionally the extension of the platform region into thelower root portion of the blade assembly simplifies the support of theplatform by using the airfoil root to clamp the platform assembly to thedisk attachment.

Turbine blades experience large radial pull loads due to centripetalacceleration. The simplest blades consist of an airfoil connecteddirectly to a root with no integral platform. This simplifies the designof a ceramic matrix composite blade. In ceramic matrix composite thedirection of the fibers control the strength of the ceramic matrixcomposite. For blades in particular, a strong biasing of the fibers inthe radial direction imparts high strength with minimal material. Theimposed loads on a blade of this type are dominated by the radial pulland for the airfoil region this creates a large radial tension stress.Bending stress in the airfoil is typically much less. The only region ofthe blade with large bending stresses is in the root region where itflares outward to form the attachment feature. In this region the ILT(Interlaminar Tensile strength, which is the layer to layer bond tensilestrength) of the ceramic matrix composite material is more limiting thatthe tensile strength. Thus minimizing ILT bending stresses throughsimplification, and/or reducing the mechanisms that create bendingstresses and/or increasing the ILT strength are desired.

Adding a platform to the existing airfoil/root assembly is challengingas the centrifugal loads created by the cantilevered platform, extendingoutward from the airfoil, create large bending stresses in the base ofthe platform where it connects with the airfoil. If the ceramic matrixcomposite platform was integrally woven into the airfoil region, thenthe fibers at the airfoil to platform intersection would be exposed toboth the large radial load of the airfoil and the local bending stressesof the cantilevered platform. Since ITL strength is much lower than thetensile strength, the design would be limited by the local bendingstresses in the platform, and not the airfoil region, resulting in lowercapability.

Separating the blade into and airfoil/root assembly and a platformassembly allows the ceramic matrix composite structure to be optimizedfor the loads imposed on those assemblies. For the airfoil/root assemblythis means a structure with minimal bending loads and a relativelystraight load path between the airfoil and the root region. For theplatform assembly a ceramic matrix composite structure with improved ITLcapability and the ability to create a platform assembly with complexfeatures is desired. Additionally, if the platform assembly was extendedto cover the attachment region of the airfoil/root assembly, the complexbending stresses in the root portion of the airfoil/root assembly couldbe distributed over a larger region of the root, and lower the magnitudeof the ILT stresses in the root, thus increasing the capability of thatregion.

ILT strength is the strength of the bond between the layers of a ceramicmatrix composite. For typical ceramic matrix composites the fiberstrength is higher than the matrix strength. Thus for typicalassemblies, created by stacking layers, the weaker matrix is thelimiting portion, and sets the design limit to the ILT stress of theassembly. In three-dimensional woven forms, additional fibers bridge thelayers, such that they increase the ILT capability. A platform assemblymade from three-dimensional ceramic matrix composite would have higherILT capability. Additionally, programmable weaving looms can createcomplex 3-D woven shapes that can accommodate thickness and dimensionalchanges through an automated process much faster than by traditionallayer by layer assembly techniques. Thus it is desirable to make aceramic matrix composite platform assembly from a 3-D woven preform forits increased ILT performance with increased complexity.

Bonding an airfoil/root assembly, with optimized ceramic matrixcomposite construction to a platform assembly with a three-dimensionalceramic matrix composite construction would create a blade assembly withthe required complexity of an integral platform and the improvedstructural performance of due to the combined construction. Additionallythe presence of the platform assembly, encasing the airfoil root region,further improves the ILT performance of the airfoil/root assembly bydistributing the attachment loads imposed by the turbine disk.

Other details of the ceramic matrix composite blade with integralplatform using a complex weave perform are set forth in the followingdetailed description and the accompanying drawings wherein likereference numerals depict like elements.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic representation of a component which can be used ina gas turbine engine;

FIG. 2 is a sectional view of the blade of FIG. 1; and

FIG. 3 is a flow chart showing the assembly method for forming the bladeof FIG. 1.

DETAILED DESCRIPTION

Referring now to the drawings, there is shown a ceramic matrix compositeblade 10 for use in a gas turbine engine (not shown). The blade 10 maybe a turbine blade used in the hot section of the engine.

The blade 10 has an airfoil portion 12 and a root portion 14. Theairfoil portion 12 and the root portion 14 may be an integral structureformed from a plurality of plies 18 of a ceramic matrix compositematerial as shown in FIG. 2.

As can be seen from FIG. 1, the blade 10 also has a platform 20 and oneor more optional buttresses 22 formed from a platform assembly structure24. The platform assembly structure 24 is formed from a ceramic matrixcomposite material. The platform assembly 24 process begins with afibrous pre-form which is in turn infiltrated with ceramic matrix toform a rigid ceramic matrix composite. The fibrous pre-form consists ofa combination of three dimensional woven structures and/or portions madefrom chopped fibers and/or two dimensional woven cloth 26. Twodimensional woven cloth typically has fiber/tow bundles interwoven suchthat a large flat sheet is created with a thickness of the sheet beingapproximately twice the thickness of the fiber/tow bundles.Three-dimensional woven preforms consists of fiber/tow bundles that arewoven in such a manner as to have additional fiber/tow bundles such thatthe thickness can be increased, and complex shapes created where localthick sections can be added and still retain connectivity to the thinsections with continuous fiber/tow connectivity. As will be discussedhereinafter, the platform assembly structure 24 is bonded to the airfoilportion 12 and root assembly 14 so as to form an integral structure.

Referring now to FIG. 3, the method of forming the blade 10 includes thestep 100 of forming the airfoil/root assembly by laying up plies 18 of aceramic matrix composite material in a mold and infiltrating the plieswith a matrix material. The plies 18 may be formed from auni-directional tape and/or a fabric or woven material such that astrong primary structure is created that can transmit the radial pull ofthe blade airfoil 12 into the root attachment region 14. A fabric may bemade from fibers called tows. Individual tows are woven together tocreate the fabric. Unidirectional-Tape can be made from a collection ofindividual fibers or a collection of tows, bonded together to form acontinuous sheet of uniform thickness. The Unidirectional tape can becut, like fabric and stacked together with plies. After the plies 18have been laid up, they may be joined together to form the airfoil/rootassembly using low temperature polymerization, high temperaturepolymerization and/or pyrolosis techniques, or bonding with a Siliconinterfacial layer.

As shown in step 102, the platform assembly structure 24 is formedseparately from the airfoil/root assembly. The platform assemblystructure 24 may be formed from a ceramic matrix composite. For example,the structure 24 may be formed using a plurality of three dimensional orchopped fibers which have been infiltrated by a matrix material. Hereagain, bonding may be accomplished using low temperature polymerization,high temperature polymerization, and/or pyrolosis, or bonding with aSilicon interfacial layer. The structure 24 may be formed in a mold.Further, the structure 24 is formed to have a central opening 30 whichextends from the top to the bottom of the structure 24. In other words,the structure 24 has a hollow core. The structure 24 may be fabricatedwith one or more chord-wise spaced apart buttresses 22. If desired, thebuttresses 22 may be omitted.

The fibers used to form the platform assembly structure 24 may includefibers such as silicon carbide, aluminum oxide, silicone nitride,carbon, and combinations thereof.

The matrix used to form the platform assembly structure and/or theairfoil/root assembly may include magnesium aluminum silicate, magnesiumbarium aluminum silicate, lithium aluminum silicate, barium strontiumaluminum silicate, bariums aluminum silicate, silicon carbide, siliconnitride, aluminum oxide, silicon aluminum oxynitride, aluminum nitride,zirconium oxide, zirconium nitride, and/or hafnium oxide.

In step 104, the airfoil/root assembly is inserted into the opening 30in the structure 24 so that the outer edge 32 of the root portion isabutted by an inner edge 34 of the structure 24.

In step 106, the platform assembly structure 24 is bonded to theairfoil/root assembly to form the blade 10. The bonding step may becarried out by introducing the matrix material and heating to densifythe ceramic matrix composite material and bond the airfoil/root assemblyto the platform assembly. The platform assembly 24 may be formed so thata portion of the platform assembly 24 may extend radially inward andcover a root region of the airfoil root assembly. Alternatively thebonding step may be carried out by introducing a bonding agent such assilicon, which after bonding creates a interfacial layer between theairfoil/root assembly and the platform assembly. Silicon, deposited in alayer on the blade/attachment assembly and/or the platform assembly,would then disperse into the resulting assembly when heated andconstrained appropriately, forming a continuous bond between theairfoil/attachment assembly and the platform assembly.

In step 108, any protruding portion, such as fibers, may be ground off.

As shown in FIG. 2, the platform root section 40 is integrated onto theblade root section 14 such that the contact between the final blade 10and a disk 42 occurs on the exterior surface 44 created by the threedimensional woven platform root section.

While the present disclosure has been described in the context offorming a turbine blade, the method could also apply to the manufactureof other components for use in a gas turbine engine.

There has been described herein a ceramic matrix composite blade withintegral platform using complex weave perform. While the ceramic matrixcomposite blade has been described in the context of specificembodiments thereof, other unforeseen alternatives, modifications, andvariations may become apparent to those skilled in the art having readthe foregoing description. Accordingly, it is intended to embrace thosealternatives, modifications and variations which fall within the broadscope of the appended claims.

What is claimed is:
 1. A method of forming a component for use in a gasturbine engine comprising the steps of: forming an airfoil/rootassembly; creating a platform assembly structure from a threedimensional woven material having an opening; inserting saidairfoil/root assembly into said opening; and bonding said platformassembly structure to said airfoil/root assembly to form said component,wherein the bonding step is carried out by introducing a bonding agent,which after bonding creates an interfacial layer between theairfoil/root assembly and the platform assembly structure.
 2. The methodof claim 1, wherein said airfoil/root assembly forming step comprisescreating said airfoil/root assembly from at least one of uni-directionaltape layers and/or fabric layers.
 3. The method of claim 1, wherein saidplatform assembly structure creating step comprises forming a portion ofthe platform assembly structure so that said portion extends radiallyinward and covers a root region of said airfoil/root assembly.
 4. Themethod of claim 1, wherein said platform assembly structure creatingstep comprises creating said platform assembly structure so that contactbetween said airfoil/root assembly and a disk occurs on an exteriorsurface of a root section of said platform assembly structure.
 5. Themethod of claim 1, wherein said platform assembly structure creatingstep comprises creating integral buttress structures located spacedapart in the chord-wise direction along the length of said platformassembly structure.
 6. The method of claim 1, wherein said bonding stepcomprises introducing a matrix material onto said platform assemblystructure and said airfoil/root assembly and heating the assemblies tobond said platform assembly structure to said airfoil/root assembly andto densify the matrix material.
 7. The method of claim 1, furthercomprising grinding off any protruding portion of said component.
 8. Themethod of claim 1, wherein the bonding step comprises depositing siliconin a layer on at least one of the airfoil/root assembly and the platformassembly structure, and then dispersing the silicon into the resultingassembly when heated and constrained, thereby forming a continuous bondbetween the airfoil/root assembly and the platform assembly structure.